On account of the high turbine inlet temperatures which are normal nowadays in modern gas turbines or gas turbosets, at least the blades of the first turbine stages must often be cooled.
To cool the blades, a cooling fluid, frequently air, which has been extracted from the compressor, is passed through flow passages inside the airfoil. The cooling is effected by convective heat transfer from the passage walls to the cooling fluid. The heated cooling fluid is then often delivered through bores in the region of the trailing edge of the blade into the blade surroundings. A turbine blade fluidically cooled in this way has been disclosed, for example, by U.S. Pat. No. 4,820,123 or also by European Patent Application EP 0 649 975 A1.
On account of the ever increasing turbine inlet temperatures of modern gas turbines, it is also often necessary to cool all the components which are subjected to the flow of hot gas. Therefore, not only the airfoil but also the blade root has to be cooled. Furthermore, in particular blades of stators are often provided with shroud elements, which are then likewise to be at least partly cooled. A conventional method for cooling the shroud elements is for some of the cooling fluid which flows through the airfoil for cooling the latter to be passed through a cooling-fluid bore provided in the shroud element and for it to then be delivered outward into the surroundings of the blade. From here, the released cooling fluid ultimately passes via component gaps into the main flow of the turbine. However, due to the inflow of the cooling fluid via component gaps, flow losses of the main flow are caused on the one hand. Furthermore, the delivered cooling fluid is often still not thermally consumed for cooling purposes, so that this also results in a thermodynamic loss. A higher cooling-fluid mass flow is required for a required cooling capacity. Both these factors lead to a deterioration in the efficiency of the gas turbine or gas turboset. Cooling of the blades which is not adapted may also lead to a reduced service life of the blades.
In the case of fluidically cooled shroud elements known from the prior art, uneven distribution of the cooling capacity is also often to be observed. This often affects in particular the corner regions of the shroud elements. This may lead to the corner regions of the shroud elements not being adequately cooled and in this case to “hot spots”.
On the other hand, uniform distribution of the cooling capacity over the shroud element is also often not required, but rather there are regions which have to be cooled to a greater extent, and other regions which necessitate only a lower degree of cooling. The cooling effect of the cooling fluid is therefore often not optimal.
U.S. Pat. No. 5,320,485 discloses a turbomachine blade, in particular a turbine blade of a gas turbine or gas turboset, having a blade root, a blade tip and an airfoil which extends between the blade root and the blade tip in a blade longitudinal direction and has a suction side and a pressure side, a shroud element being arranged on the blade tip.